(a) Field of the Invention
The present invention relates generally to spacecraft or satellite attitude determination systems and, more particularly, to a method and apparatus for correcting gyro scale factor and misalignment errors to improve attitude determination performance.
(b) Description of Related Art
The term attitude is used to describe the orientation of an object with respect to a reference orientation. Attitude is of particular interest in satellite or spacecraft operations. For example, if a satellite is used in a communications application, it is necessary that the satellite be oriented in the proper direction to receive and/or transmit relevant information for the communication link.
The attitude of a satellite is determined by computations based on the output of sensors located on the satellite. Gyros and object trackers (such as star trackers, sun sensors, and earth sensors) are two types of sensors that may be used in attitude determination systems. In general, gyros are used to measure the rate at which the spacecraft is moving. By integrating gyro output, spacecraft attitude may be determined. The use of gyros in conjunction with star trackers is commonly known in the art as a stellar-inertial attitude determination system.
Object trackers such as star trackers, earth sensors, or sun sensors are used to determine the orientation or attitude of the satellite with respect to the objects being tracked. Object trackers commonly use a CCD array to measure heat or light emitted from the tracked object. Typically, satellites use an ephemeris system to determine the location of the objects being tracked with respect to the earth. An ephemeris system uses a table containing the coordinates of a celestial body at a number of specific times during a given period. Using ephemeris techniques and object trackers, spacecraft attitude with respect to the earth can be determined.
The use of gyros in an attitude determination system results in attitude errors induced by gyro scale factor and misalignment errors (GSFME). These errors appear as gyro bias errors for non-dynamic spacecraft missions such as an orbit-normal steered spacecraft orbiting in a circular orbit. In non-dynamic missions, the angular accelerations of the spacecraft about its axes are relatively small. Thus, bias errors in non-dynamic missions can easily be estimated and compensated for in the spacecraft's attitude determination system by using a standard six state stellar/inertial Kalman filter, which is well known in the art. Therefore, GSFME has not typically been of much concern because most spacecraft missions were historically not dynamic.
However, spacecraft use has expanded to dynamic missions such as sunnadir steering, perigee passing for highly elliptical orbit (HEO), and dynamic target tracking. In dynamic missions, the angular accelerations of the spacecraft about its axes are relatively large. Thus, GSFME in dynamic spacecraft missions can contribute significantly to spacecraft attitude error. Moreover, in dynamic missions, GSFMEs do not manifest themselves as bias and, therefore, cannot be easily eliminated using a six-state Kalman filter.
Accordingly, there is a need for a method and apparatus that eliminates gyro misalignment and bias errors in dynamic spacecraft missions.